Method of Repairing, Splicing, Joining, Machining, and Stabilizing Honeycomb Core Using Pourable Structural Foam and a Structure Incorporating the Same

ABSTRACT

A method of the present disclosure includes of repairing a core stiffened structure with structural foam. Another method includes splicing core members together using structural foam. Another method includes joining a core member to a structure using structural foam. Another method includes using structural foam to stabilize a core member during a machining process. Another method includes stabilizing a core member with structural foam to prevent the core member from crushing in autoclave pressure. The present disclosure further includes a core stiffened structure have a core member with structural foam therein.

BACKGROUND

1. Technical Field

One method of the present disclosure relates to using a pourablestructural foam in the repair of a honeycomb core stiffened structure.Another method of the present disclosure relates to the splicing ofhoneycomb core using pourable structural foam. Another method of thepresent disclosure relates to the joining of honeycomb core to astructural with the use of structural foam. Another method of thepresent disclosure relates to using structural foam to provide stabilityto honeycomb core during a machining process. One apparatus of thepresent disclosure includes a core stiffened structure. Another methodof the present disclosure relates to the utilization of structural foamfor the stabilizing of honeycomb core during an autoclave curing cycle.

2. Description of Related Art

In a conventional repair process, a damaged core stiffened structure isrepaired by removal/replacement of the damaged honeycomb core with newhoneycomb core. Such a process can be expensive, time-consuming, and canrequire multiple curing and bonding processes. There is a need for animproved method of repairing a honeycomb core stiffened structure.

In a conventional honeycomb core splicing process, exposed cell wallsare meticulously bonded to adjacent exposed cell walls of anotherhoneycomb core with an adhesive. Such a process can be expensive,time-consuming, and can require an adhesive cure cycle. Further, such aconventional process can be inadequate for splicing honeycomb coremembers having non-compatible materials. There is a need for an improvedmethod of splicing portions of honeycomb core.

In a conventional honeycomb core joining process, exposed cell walls arejoined to adjacent structure with an adhesive. Such a process can beexpensive, time-consuming, and require an adhesive cure cycle. There isa need for an improved method of joining honeycomb core to adjacentstructure.

In a conventional honeycomb core machining process, the honeycomb coreis machined to a desired contour at a feed rate low enough to preventdamage to the unstable honeycomb core. Such a process can betime-consuming. There is a need for an improved method of machininghoneycomb core, especially for large cell carbon prepreg core.

In a conventional curing cycle of the honeycomb core stiffenedstructure, the autoclave pressure can cause the honeycomb core to crushunless a low bevel angle on the honeycomb core is used. Conventionally,if the geometry of the core stiffened panel did not allow for the lowangle bevel, then the core stiffened panel would have to bemanufacturing using labor intensive and procedures (multiple curecycles) that can also add significant weight penalties. Hence there is aneed for improved method of stabilizing honeycomb core during anautoclave cure cycle.

DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the methods andapparatuses of the present disclosure are set forth in the appendedclaims. However, each method and apparatus, as well as a preferred modeof use, and further objectives and advantages thereof, will best beunderstood by reference to the following detailed description when readin conjunction with the accompanying drawings, wherein:

FIG. 1 is a side view of a illustrative rotorcraft;

FIG. 2 is a top view of a rotor blade, according to an illustrativeembodiment;

FIG. 3 is a cross-sectional view of the rotor blade, taken from sectionlines 3-3 in FIG. 2;

FIG. 4 is a schematic view of a method of repairing a honeycomb corestiffened structure, according to an illustrative embodiment;

FIG. 5 is a top view of a portion of a rotor blade, for illustrativepurposes;

FIG. 6 is a top view of a portion of a rotor blade, for illustrativepurposes;

FIG. 7 is a top view of a portion of a rotor blade, for illustrativepurposes;

FIG. 8 is a stylized view of components of the foam system, according toan illustrative embodiment of present application;

FIG. 9 is a front view of a mixer used to mix the foam system, accordingto an illustrative embodiment of present application;

FIG. 10 is a stylized view of a mixture of the foam system, according toan illustrative embodiment of present application;

FIG. 11A is a cross-sectional view of the rotor blade, taken fromsection lines 11A-11A in FIG. 7;

FIG. 11B is a cross-sectional view of the rotor blade, similar to FIG.11A;

FIG. 12 is a schematic view of a method of splicing together one or moresections of core material, according to an illustrative embodiment;

FIG. 13 is a top stylized view of a method of splicing together one ormore sections of core material, according to an illustrative embodiment;

FIG. 14 is a top stylized view of a method of splicing together one ormore sections of core material, according to an illustrative embodiment;

FIG. 15 is a top stylized view of a method of splicing together one ormore sections of core material, according to an illustrative embodiment;

FIG. 16 is a top stylized view of a method of splicing together one ormore sections of core material, according to an illustrative embodiment;

FIG. 17 is a schematic view of a method of joining together corematerial to a structure, according to an illustrative embodiment;

FIG. 18 is a top stylized view of a method of joining together corematerial to a structure, according to an illustrative embodiment;

FIG. 19 is a schematic view of a method of machining honeycomb core,according to an illustrative embodiment;

FIG. 20 is a cross-sectional stylized view of a method of machininghoneycomb core, according to an illustrative embodiment;

FIG. 21 is a cross-sectional stylized view of a machined honeycomb core,according to an illustrative embodiment;

FIG. 22 is a top view of a core stiffened structure, according to anillustrative embodiment;

FIG. 23 is a front view of a core stiffened structure, according to anillustrative embodiment;

FIG. 24 is a cross-sectional view of a core stiffened structure, takenat section lines 24-24 in FIG. 22, according to an illustrativeembodiment;

FIG. 25 is a cross-sectional view of a core stiffened structure,according to an illustrative embodiment;

FIG. 26 is a cross-sectional view of a core stiffened structure,according to an illustrative embodiment;

FIG. 27 is a cross-sectional view of a core stiffened structure,according to an illustrative embodiment;

FIG. 28 is a cross-sectional view of a core stiffened structure,according to an illustrative embodiment;

FIG. 29 is a schematic view of a method of stabilizing honeycomb corefor an autoclave cure cycle of a core stiffened structure, according toan illustrative embodiment;

FIG. 30 is a top view of a core stiffened structure, according to anillustrative embodiment;

FIG. 31 is a front view of a core stiffened structure, according to anillustrative embodiment;

FIG. 32 is a cross-sectional view of a core stiffened structure, takenat section lines 32-32 in FIG. 30, according to an illustrativeembodiment;

FIG. 33 is a cross-sectional stylized view of a method of stabilizinghoneycomb core for an autoclave cure cycle of a core stiffenedstructure, according to an illustrative embodiment; and

FIG. 34 is a cross-sectional stylized view of a method of stabilizinghoneycomb core for an autoclave cure cycle of a core stiffenedstructure, according to an illustrative embodiment.

DETAILED DESCRIPTION

Illustrative embodiments of the methods and apparatuses are describedbelow. In the interest of clarity, all features of an actualimplementation may not be described in this specification. It will ofcourse be appreciated that in the development of any such actualembodiment, numerous implementation-specific decisions must be made toachieve the developer's specific goals, such as compliance withsystem-related and business-related constraints, which will vary fromone implementation to another. Moreover, it will be appreciated thatsuch a development effort might be complex and time-consuming but wouldnevertheless be a routine undertaking for those of ordinary skill in theart having the benefit of this disclosure.

In the specification, reference may be made to the spatial relationshipsbetween various components and to the spatial orientation of variousaspects of components as the devices are depicted in the attacheddrawings. However, as will be recognized by those skilled in the artafter a complete reading of the present disclosure, the devices,members, apparatuses, etc. described herein may be positioned in anydesired orientation. Thus, the use of terms such as “above,” “below,”“upper,” “lower,” or other like terms to describe a spatial relationshipbetween various components or to describe the spatial orientation ofaspects of such components should be understood to describe a relativerelationship between the components or a spatial orientation of aspectsof such components, respectively, as the device described herein may beoriented in any desired direction.

Referring now to FIG. 1 in the drawings, a rotorcraft 101 isillustrated. Rotorcraft 101 has a rotor system 103 with a plurality ofrotor blades 205. The pitch of each rotor blade 205 can be manipulatedin order to selectively control direction, thrust, and lift ofrotorcraft 101. Rotorcraft 101 can further include a fuselage 107,anti-torque system 109, and an empennage 111. A variety of members onrotorcraft 101 can be manufacturing as a core (honeycomb, foam, orhoneycomb/foam) stiffened structure, as discussed further herein.

Referring now also to FIGS. 2 and 3, rotor blade 205 is an example of arotorcraft member than can be configured with core to efficientlyprovide strength and stiffness. Rotor blade 205 can include a leadingedge portion 207, a trailing edge portion 209, a root end portion 203,and a tip end portion 205. In the illustrated embodiment, an upper skin211 and a lower skin 213 form an aerodynamic surface with a spar member215 and a core member 217 forming the internal structure.

It should be appreciated that rotorcraft 101 is merely illustrative ofan aircraft for with which one or more methods of the present disclosurecan be utilized with regard to. Further, rotor blade 205 is merelyillustrative of the wide variety of core stiffened structures that canutilize one or methods disclosed herein. In another embodiment, ahorizontal stabilizer 115 is the core stiffened structure, for example.

Referring now to FIG. 4, a method 401 of repairing a honeycomb corestiffened structure is schematically illustrated. During operation ofrotorcraft 101, damage to a core stiffened structure can occur in avariety of scenarios, such as a collision with an object. In order toensure safety of the rotorcraft, the damaged portion of the corestiffened structure must be repaired. It should be appreciated that thedamage can exist in the skin and/or the honeycomb core. However, removalof the damaged skin may also require repair of the honeycomb core sincethe skin and the honeycomb core are bonded together. For illustrativepurposes, method 401 is illustrated and described herein with regard torepair of rotor blade 205; however, it should be appreciated that method401 can be performed on any core stiffened structure.

Method 401 can include a step 403 of removing the damaged portion ofskin, such as upper skin 211, and the damaged portion of honeycomb core,such as core member 217. For illustrative purposes, FIG. 5 shows thedamaged skin portion removed by a cutout, which in the illustratedembodiment is circular. Further, FIG. 6 shows the damaged core portionremoved, which creates a cavity 219. Removal of the damaged skin portionand the damaged core portion can be performed with a circular saw, forexample.

Method 401 can further include a step 405 for cleaning and removing anycontaminants in and around cavity 219. Step 405 can include using acleaning solution, such as an alcohol based fluid to removecontaminants.

Method 401 can include a step 407 for mixing and preparing a foammixture. Referring now also to FIGS. 8-10, the foam system used inmethod 401 is preferably a pourable urethane foam system that expands tofill cavity 219 of core member 217. Upon curing, the foam is preferablyrigid and flexible, while having a relatively low density, or a densitysimilar to the core. In the illustrated embodiment, the foam system is apolyisocyanurate pour foam system marketed under the name of StepanfoamBX 450, by the Stepan Chemical Company. It should be fully appreciatedthat other foam systems may be used. The Stepanfoam BX 450 includes twoseparate resin components, namely a resin T (70% by weight) 805 and aresin R (30% by weight) 803. The desired weight amounts of resin T 805and resin R 803 are measured and poured into a container 801. The resinT 805 and resin R 803 are mechanically agitated to form a homogenousmixture 807. A mixer 901 can be used in rotation in order to mix resin T805 and resin R 803 to form mixture 807. Mixer 901 includes a shaft 903,the shaft 903 being configured to be driven by a drill motor, or thelike. Mixer 901 may include a plurality of blades located on a firstblade disc 905 and a second blade disc 907. In the illustratedembodiment, mixer 901 is rotated at 3500 revolutions per minute for10-15 seconds when mixing resin T 805 and resin R 803 to form mixture807.

A step 409 includes pouring foam mixture 807 into cavity 219. Once themixing in step 409 is complete, it is preferred that mixture 807 ispoured into cavity 219 as quickly as possible. Once foam mixture 807 ispoured into cavity 219, foam mixture 807 is allowed to expand and fillthe space of cavity 219 so as to form an expanded foam 221 (shown inFIG. 7). In the illustrated embodiment, the foam mixture 807 is allowedto “free rise”; however, an alternative embodiment can includerestricting the expansion by placing an upper mold over the cavity 219.Preferably, the upper mold would include an opening that allows therelease of outgases.

A step 411 includes trimming the expanded foam 221 after foam mixture807 has fully expanded and become semi-rigid. Trimming the expanded foam221 includes reducing the outer surface to correspond approximately withthe outer airfoil surface of upper skin 211. Step 411 can be performedby any variety of processes, such as a cutting, sanding, machining, toname a few examples. FIG. 11A is an illustration of expanded foam 221trimmed such that the outer surface corresponds approximately to theouter airfoil surface of upper skin 211.

A step 413 includes laying up a composite skin patch 223 on the expandedfoam 221 and allowing the repair assembly to cure. An illustrativecomposite skin patch 223 is shown in FIG. 11B. In the illustratedembodiment, composite skin patch 223 is a plurality of strips ofcomposite fibers with resin (i.e. pre-preg). Composite skin patch 223can be laid directly onto the outer surface of expanded foam 221.Composite skin patch 223 and expanded foam 221 are allowed to cure. Inthe illustrated embodiment, the curing of composite skin patch 223 andexpanded foam 221 takes place at room temperature, such that anartificial heat source is not required. In another embodiment, skinpatch 223 can be a precured composite skin patch or of another rigidmaterial, such as a metal skin patch. In such an embodiment, the rigidskin patch can be adhesively bonded to the expanded foam 221 with anadhesive.

Method 401, as described further herein, includes significant advantagesof conventional repair methods. For example, method 401 does not requirespecial repair tooling that would otherwise be required to separatelylay-up and cure a composite patch. Further, method 401 does not requirethermal equipment that would be required if an adhesive as used thatrequired an elevated temperature to cure. Further, method 401 can beperformed in the field such that repairs to composite core stiffenedstructures do not have to be performed in a facility having specializedequipment. Further, when method 401 is performed on composite corestiffened structure wherein the composite core is large cell core (cellsgreater than 0.5 inches), the expanded foam 221 (as trimmed) acts as atooled surface so as to provide uniform support of the composite skinpatch 223 prior and during curing.

It should be appreciated that method 401 can be performed on a widevariety of core-stiffened structures. Further, the core-stiffenedstructures can have a core member 217 comprising any variety ofmaterials, such as carbon, paper, Nomex®, fiberglass, Kevlar®, plastic,metal, foam, to name a few examples. Further, core member 217 can havecell members of a variety of shapes, such as square, hexagonal,elliptical, round, to name a few examples. Further, core member 217 maynot even have cell members, particularly when core member 217 is a foammaterial.

Referring now to FIG. 12, a method 1201 of splicing together one or moresections of core material is schematically illustrated. Duringfabrication of a core stiffened structure, such as rotor blade 205, itcan be necessary to splice one or more sections of core materialtogether. However, conventional methods of splicing core material havesignificant deficiencies, as described further herein.

Referring also to FIG. 13, method 1201 can include a step 1203 ofpositioning a first core member 1301 adjacent to a second core member1303. Step 1203 can be implemented in a variety of environments. Forexample, tooling can be used for aligning first core member 1301adjacent to second core member 1303; however, tooling is not necessarilyrequired. One advantage of using foam for splicing first core member1301 to second core member 1303 is the increase in bonding surfacereduces the importance of bond surface alignment that can exist in aconventional adhesive bonding procedure.

Step 1205 includes preparing and mixing the foam mixture, which issimilar to step 407 described herein with regard to method 401. Method1201 also includes a step 1207 for pouring the foam mixture into thesplicing area common to the cell walls of first core member 1301 andsecond core member 1303. In another embodiment, step 1207 isaccomplished by pouring the foam mixture into a cavity below splicingarea so that the foam mixture expands and rises into the splicing area,similar to as described with regard to FIG. 33. Such a technique may beparticularly desirable when the cells are relatively small and/or thecell walls lack porosity. The foam mixture is allowed to expand into anexpanded foam. A step 1209 includes curing the expanded foam. In theillustrated embodiment, the expanded foam can be cured at roomtemperature in a relatively short about of time. One advantage of method1201 is that the process does not require thermal equipment that mayotherwise be required to cure a conventional adhesive. A step 1211 caninclude trimming the expanded foam to be flush with the upper and lowersurfaces of first core member 1301 and second core member 1303. Step1211 can also include machining the first core member 1301 and thesecond core member 1303 to a desired contour while also trimming theexpanded foam.

Referring now also to FIGS. 14-16, example embodiments of core membersspliced together using method 1201 are illustrated. Referring inparticular to FIG. 14, first core member 1301 and second core member1303 are both honeycomb composite core members of the same size andmaterial. Expanded foam 1305 is bonding contact with the adjacent cellwalls of first core member 1301 and second core member 1303. Referringin particular to FIG. 15, first core member 1301 and second core member1303 are in bonding contact with expanded foam 1305. In the illustratedembodiment, second core member 1303 has smaller sized cell memberscompared to first core member 1301. The illustrated embodiment is meantto demonstrate how the expanded foam 1205 is particularly advantageousfor splicing together core members having different geometry and/or cellsizes. As illustrated, expanded foam 1205 provides bonding contact withthe cell walls of first core member 1301 and second core member 1303even though the geometry of the cell walls don't provide symmetricbonding surfaces. Further, the splicing of two different core members,each having different geometries, can be particularly useful whentailoring stiffness across a structure. For example, rotor blade 205 canhave spliced core members such that the larger core member is locatedmore towards spar member 215 while the smaller core member is locatedtowards the trailing edge 209. It should be appreciated that method 1201can be employed to splice any number of core members together, the coremembers having any variety of cell sizes and geometries.

Referring in particular to FIG. 16, first core member 1301 and secondcore member 1303 are in bonding contact with expanded foam 1305. In theillustrated embodiment, second core member 1303 is a foam core that doesnot have honeycomb shaped cell members. Second core member 1303 can beany type of foam of supporting structure having a homogenous materialcontinuity. For example, second core member can be of a polyvinylchloride (PVC) type material. Method 1201 is particularly well suitedfor splicing a traditional honeycomb core member to a solid core sincethe expanded foam 1305 adheres well to both members, unlike traditionalsplicing adhesives.

Referring now to FIG. 17, a method 1701 of joining together a corematerial to a structure is schematically illustrated. During fabricationof a core stiffened structure, such as rotor blade 205, it can bedesirable to abutly join a core member, such as core 217, to astructure, such as spar 215 (shown in FIG. 3). However, conventionalmethods of joining core material to structure have significantdeficiencies, as described further herein.

Referring also to FIG. 18, method 1701 can include a step 1703 ofpositioning a core member 1801 adjacent to a structure 1803. Step 1703can be implemented in a variety of environments. For example, toolingcan be used for aligning core member 1801 adjacent to structure 1803;however, tooling is not necessarily required. One advantage of usingfoam for joining core member 1801 to structure 1803 is the increase inbonding surface prevents the meticulous adhesive application alongmating portions of core member 1801 to structure 1803, which mayotherwise be required in a conventional process.

Step 1705 includes preparing and mixing the foam mixture, which issimilar to step 407 described herein with regard to method 401. Method1701 also includes a step 1707 for pouring the foam mixture into thejoining area common to the exposed cell walls of core member 1801 andthe exposed surface of structure 1803. In another embodiment, step 1707is accomplished by pouring the foam mixture into a cavity below joiningarea so that the foam mixture expands and rises into the joining area,similar to as further described with regard to FIG. 33. Such a techniquemay be particularly desirable when the joining area is relatively smalland the cell walls lack porosity. The foam mixture is allowed to expandinto an expanded foam 1805. A step 1709 includes curing expanded foam1805. In the illustrated embodiment, the expanded foam 1805 can be curedat room temperature in a relatively short amount time. One advantage ofmethod 1701 is that the process does not require thermal equipment thatmay otherwise be required to cure a conventional adhesive. A step 1711can include trimming the expanded foam 1805 to be flush with the upperand lower surfaces of core member 1801 and structure 1803.

Referring now to FIG. 19, a method 1901 of machining honeycomb core isschematically illustrated. During fabrication of a core stiffenedstructure, such as rotor blade 205, it is necessary to machine thehoneycomb core member, such as core 217 (shown in FIG. 3), to a desiredcontour. However, conventional methods of machining core material havesignificant deficiencies, as described further herein.

Referring also to FIGS. 20 and 21, method 1901 can include a step 1903of preparing and mixing the foam mixture, which is similar to step 407described herein with regard to method 401. Method 1901 also includes astep 1905 of pouring the foam mixture into the cells of a stock coremember 2001. In another embodiment, step 1907 is accomplished by pouringthe foam mixture into a cavity below the cells so that the foam mixtureexpands and rises into the cells, similar to as further described withregard to FIG. 33. Such a technique may be particularly desirable whenthe cell are relatively small and/or the cell walls lack porosity. Thefoam mixture is allowed to expand into an expanded foam 2003. A step1907 includes curing expanded foam 2003. In the illustrated embodiment,the expanded foam 2003 can be cured at room temperature.

Method 1901 includes a step 1909 of machining the stock core member 2001containing expanded foam 2003 to a desired contour. A contouring machine2005 can be used to controllable machine away undesired portions ofstock core member 2001 containing expanding foam 2003 until the desiredcontour is achieved. Referring to FIG. 21, a final foam filled coremember 2007 is illustrated. Using expanded foam 2003 to support thehoneycomb core during machining not only increases the speed at whichthe honeycomb core can be machined, but also allows the honeycomb coreto be machined at more severe angles, smaller thicknesses, and complexcontours. Further, expanded foam 2003 adds strength and stiffness tostock core member 2001.

Referring now to FIGS. 22-24, a core stiffened structure 2201 that ismanufactured utilizing one or more methods described herein isillustrated. Structure 2201 can be any of a variety of structuralmembers. For example, structure 2201 can be a panel 113 on rotorcraft101, as illustrated in FIG. 1. Structure 2201 can have an upper skin2203, a lower skin 2205, and a honeycomb core 2207 filled with anexpanded foam 2209. Honeycomb core 2207 can be filled with expanded foam2209 and machined with method 1901, as described further herein.

Structure 2201 is illustrative of one unique advantage of utilizinghoneycomb core 2207 filled with expanded foam 2209. Not only can thehoneycomb core be machined faster and more accurately when filled withexpanded foam, but structure 2201 can be manufactured so that thestructure itself is more efficient. Unlike traditional composite corestructures, the end portions of honeycomb 2207 can be vertical since theexpanded foam 2209 provides the necessary strength to withstandautoclave pressure during the curing of structure 2201. In theillustrated embodiment, the end portion or ramp angle A1 is 90°;however, alternative embodiments can have a ramp angle A1 that is lessthan 90°. For example, ramp angle A1 can be 75°. Conventional compositecore structures are restricted to a ramp angle less than 27° so that thehoneycomb core does not crush during the autoclave pressure of thecuring cycle.

Referring now to FIGS. 25 and 26, a core stiffened structure 2501 thatis manufactured utilizing one or more methods described herein isillustrated. Structure 2501 can have any of a variety of structuralimplementations. For example, structure 2501 can be a floor panel withinfuselage 107 on rotorcraft 101, as illustrated in FIG. 1. Structure 2501can have an upper skin 2503, a lower skin 2505, and a honeycomb core2507 filled with an expanded foam 2509. Honeycomb core 2507 can befilled with expanded foam 2509 and machined with method 1901, asdescribed further herein.

Structure 2501 is illustrative of one unique advantage of utilizinghoneycomb core 2507 filled with expanded foam 2509. Not only can thehoneycomb core 2507 be machined faster and more accurately when filledwith expanded foam 2509, but structure 2501 can be utilized to supportout-of-plane loads with an attachment assembly, such as sleeve assembly2511, without having to use a heavy potting compound since the expandedfoam 2509 provides the necessary strength to support the loading aroundthe sleeve assembly 2511. In the illustrated embodiment, the attachmentassembly is illustrated as sleeve assembly 2511; however, it should beappreciated that the attachment assembly can be any of a variety ofattachment members configured for coupling to an apparatus. For example,sleeve assembly 2511 is particularly well suited as a mounting apparatusfor an occupant seat on the floor member 2501 of rotorcraft 101. Sleeveassembly 2511 can have an upper sleeve 2513 and a lower sleeve 2515 thatmatingly fit together. In one embodiment, sleeve assembly 2511 hasinternal threads that are configured for receiving a seat attachmentfitting.

Referring in particular to FIG. 26, structure 2601 is substantiallysimilar to structure 2501 except that expanded foam 2509 does notcompletely fill honeycomb core 2507, but rather expanded foam 2509 isparticularly located in regions where each sleeve assembly 2511 is to beinstalled. A large floor panel structure can have any plurality ofregions of expanded foam 2509 so that a desired seating configurationcan be mounted to the structure with the implementation specificattachment assembly.

Referring now to FIG. 27, a core stiffened structure 2701 that ismanufactured utilizing one or more methods described herein isillustrated. Structure 2701 can have any of a variety of structuralimplementations. For example, structure 2701 can be a horizontalstabilizer 115 on rotorcraft 101, as illustrated in FIG. 1. Structure2701 can have an upper skin 2703, a lower skin 2705, an inner support2707, an upper core member 2709, a lower core member 2711, a forwardfoam member 2713, and an aft foam member 2715. In the illustratedembodiment, upper core member 2709 and lower core member 2711 are notfilled with foam, but in an alternative embodiment upper core member2709 and a lower core member 2711 are filled with expanded foam, such asexpanded foam 2509, illustrated in FIG. 25 and further described herein.In another embodiment, only upper core member 2709 is filled with foamin order to provide increased strength and stiffness to withstand heatinduced stressed near the upper surface, the heat being discussedfurther below.

Core stiffened structure 2701 utilizes forward foam member 2713 forstructurally joining forward edges of upper core member 2709 and lowercore member 2711 together with upper skin 2703, lower skin 2705, andinner support 2707. Similarly, aft foam member 2715 structurally joinsaft edges of upper core member 2709 and lower core member 2711 togetherwith upper skin 2703, lower skin 2705, and inner support 2707. Thefilling and adhesion qualities make expandable foam uniquely capable ofjoining honeycomb core members that are out-of-plane to each other. Inother words, expandable foam can be used to join honeycomb core membersthat are not adjacent to each other in the same plane.

In the illustrated embodiment, core stiffened structure 2701 can bemanufactured by assembling the components, then subsequently filling therespective cavities to form forward foam member 2713 and aft foam member2715 with foam. Forward foam member 2713 and aft foam member 2715 can beformed by mixing pourable foam as described in step 407 of method 401,and then pouring the foam into the cavities as similarly described instep 409 of method 401. The expanded foam can be trimmed and cured, asalso further described herein.

In the illustrated embodiment, core stiffened structure 2701 is ahorizontal stabilizer which can be subjected to heat from the exhaustsystem. The dynamic heat flow typically subjects the upper surface ofthe structure 2701 to the most severe heat. In order to withstand, upperskin 2703 is formed from a composite system having a high temperatureresistance resin, such as Bismaleimide (BMI) resin. However, BMI resinis typically expensive and expensive to use. As such, the uniqueconfiguration of core stiffened structure 2701 inhibits heat transferfrom upper skin 207 to lower skin 2705 because the honeycomb core doesnot extend from the upper surface to the lower surface, rather innersupport 2707 separates the honeycomb core between upper core member 2709and lower core member 2711, thereby acting as an insulator. Further,forward foam member 2713 and aft foam member 2715 structural join uppercore member 2709 and lower core member 2711 while also inhibiting heattransfer therebetween. Such a configuration allows lower skin 2705 to bemanufactured out of a composite system having a typical non-hightemperature resin, thereby saving the expense associated with using BMIresin.

It should be appreciated that even though internal member 2707 isillustrated as an enclosed member, the geometry of internal member 2707is implementation specific. For example, an alternative embodiment ofinternal member 2707 can be C-shaped, I-shaped, or even be multiplesingle surface members, to name a few examples.

Referring now to FIG. 28, a core stiffened structure 2801 that ismanufactured utilizing one or more methods described herein isillustrated. Structure 2801 can have any of a variety of structuralimplementations. For example, structure 2801 can be a tailboom member ona rotorcraft. Structure 2801 can have an outer skin 2803, an inner skin2805, a core member 2807, and a foam member 2811. In the illustratedembodiment, core member 2807 is filled with expanded foam 2809; however,in an alternative embodiment core member 2807 is void of expanded foam2809. Expanded foam 2809 inhibits heat transfer from outer skin 2803 toinner skin 2805, as discussed further below.

Core stiffened structure 2801 utilizes foam member 2811 for structurallyjoining edges of core member 2807. The filling and adhesion qualitiesmake expandable foam uniquely capable of joining honeycomb core membersthat are out-of-plane to each other. In other words, expandable foam canbe used to join honeycomb core members that are not adjacent to eachother in the same plane.

In the illustrated embodiment, core stiffened structure 2801 can bemanufactured by assembling the components, then subsequently filling therespective cavity with foam to form foam member 2811. Foam member 2811can be formed by mixing pourable foam as described in step 407 of method401, and then pouring the foam into the cavity as similarly described instep 409 of method 401. The expanded foam can be trimmed and cured, asalso further described herein. Core member 2807 can be formed by method1901, for example.

In the illustrated embodiment, core stiffened structure 2801 is atailboom member which can be subjected to heat from the exhaust systemof the rotorcraft. The heat flow typically subjects the upper surface ofthe structure 2801 to the most severe heat. In order to withstand, anupper portion of outer skin 2803 is formed from a composite systemhaving a high temperature resistance resin, such as Bismaleimide (BMI)resin. However, BMI resin is typically expensive and expensive to use.As such, the unique configuration of core stiffened structure 2801inhibits heat transfer from outer skin 2803 to inner skin 2805 becausethe expanded foam 2809 inhibits heat flow through honeycomb core 2807,thereby acting as an insulator. Such a configuration also allows theremaining portion of outer skin 2803 and the inner skin 2805 to bemanufactured out of a composite system having a typical non-hightemperature resin, thereby saving the expense associated with using BMIresin.

Referring now to FIG. 29, a method 2901 of stabilizing honeycomb coreduring an autoclave cure of a core stiffened panel is schematicallyillustrated. Conventionally, honeycomb core is required to have lowangle beveled edge portions to prevent a crushing of the honeycomb coredue autoclave pressure in the curing cycle. Conventionally, if thegeometry of the core stiffened panel did not allow for the low anglebevel, then the core stiffened panel would have to be manufacturingusing labor intensive and procedures (multiple cure cycles) that canalso add significant weight penalties. In contrast, method 2901 of thepresent disclosure prevents the autoclave crushing of honeycomb core atany core bevel angle, while allowing the core stiffened structure to becured in a single cure cycle.

Method 2901 includes a step 2903 of mixing and preparing the foammixture. Step 2903 is substantially similar to step 407 described hereinwith regard to method 401. A step 2905 includes expanding the foammixture in the desired honey comb cells. Step 2907 includes trimming theexpanded foam. A step 2909 includes assembling the composite skins withthe honeycomb core member and curing in a cure cycle.

Referring now also to FIGS. 30-34, method 2901 is further described withregard to the manufacturing of core stiffened panel 3001. Core stiffenedpanel 3001 is illustrative of any variety of composite structures. Forexample, core stiffened panel 3001 can be a floor panel within fuselage107 of rotorcraft 101. Core stiffened panel 3001 can include an uppercomposite skin 3003, a lower composite skin 3005, and a honeycomb coremember 3007. The edge bevel portions 3011 of honeycomb core member 3007are filled with expanded foam 3009 in order to provide the strength andstiffness to prevent deformation during the autoclave cure cycle. Edgebevel portions 3011 can be of any bevel angle B1 and are not limited toa bevel angle less than 27°. In the illustrated embodiment, bevel angleB1 is approximately 60°.

Step 2905 of method 2901 includes expanding the foam mixture in thedesired honeycomb cells. Referring in particular to FIG. 33, step 2905can be accomplished by pouring the foam mixture into a cavity tool 3301,then placing the stock honeycomb core on the cavity tool 3303 so thatthe foam mixture expands and fills the desired honeycomb cells that willbe trimmed to the form the edge bevel portions 3011. In anotherembodiment, the foam mixture is poured into the cells. It should beappreciated that the edge bevel portions 3011 can be machined into thestock honeycomb core either before or after the foam mixture is expandedinto the cells.

Step 2905 can also include tailoring the density of the expanded foam3009 within honeycomb core member 3007. For example, the density of theexpanded foam tends to decrease as a function of vertical height, assuch, the height H1 of cavity tool 3301 can be chosen such that thedensity of the expanded foam 3009 within honeycomb core member 3007 iswithin a desired density range. By tailoring the density of expandedfoam 3009, the strength and stiffness can be tailoring based upon avariety of factors, such as autoclave pressure, edge bevel angle B1,honeycomb core material, to name a few. For example, it may be desirableto use a higher density of expanded foam where the edge bevel angle B1is larger than an area where the edge bevel angle B1 is smaller.

Step 2907 can include trimming the expanded foam 3009. Referring to FIG.34, a fully trimmed honeycomb core member 3007 is illustrated. In theillustrated embodiment, the expanded foam 3009 is trimmed when the edgebevel portions 3011 are machined into the stock honeycomb core.

A step 2909 includes assembling the composite skins with the honeycombcore member and curing in a cure cycle. In the illustrated embodiment,upper skin 3003 and lower skin 3305 are cured after assembly in a singlecure cycle involving autoclave pressure and heat.

The particular embodiments disclosed above are illustrative only, as theapparatuses and methods may be modified and practiced in different butequivalent manners apparent to those skilled in the art having thebenefit of the teachings herein. Modifications, additions, or omissionsmay be made to the apparatuses described herein without departing fromthe scope of the invention. The components of the system may beintegrated or separated. Moreover, the operations of the system may beperformed by more, fewer, or other components.

Furthermore, no limitations are intended to the details of constructionor design herein shown, other than as described in the claims below. Itis therefore evident that the particular embodiments disclosed above maybe altered or modified and all such variations are considered within thescope and spirit of the application. Accordingly, the protection soughtherein is as set forth in the claims below.

To aid the Patent Office, and any readers of any patent issued on thisapplication in interpreting the claims appended hereto, applicants wishto note that they do not intend any of the appended claims to invokeparagraph 6 of 35 U.S.C. §112 as it exists on the date of filing hereofunless the words “means for” or “step for” are explicitly used in theparticular claim.

1. A method of repairing a core stiffened structure, the method comprising: removing a damaged portion of the core stiffened structure; pouring a foam mixture into a cavity, the cavity being created by the step of removing the damaged portion of the core stiffened structure; allowing the foam mixture to expand and fill the cavity so as to form an expanded foam; trimming the expanded foam; and placing a skin patch on the expanded foam.
 2. The method according to claim 1, wherein the damaged portion includes a portion of a honeycomb core configured for providing stiffness to the core stiffened structure.
 3. The method according to claim 1, wherein the damaged portion includes a portion of a foam core configured for providing stiffness to the core stiffened structure.
 4. The method according to claim 1, wherein the step of removing the damaged portion of core stiffened structure comprises using a hole saw to create a circular shaped cavity in the core stiffened structure.
 5. The method according to claim 1, wherein the step of trimming the expanded foam comprises sanding an upper surface of the expanded foam to a desired contour.
 6. The method according to claim 1, wherein the step of trimming the expanded foam comprises reducing the upper surface of the expanded foam until the upper surface of the expanded foam is similar to a skin surface of the core stiffened structure.
 7. The method according to claim 1, wherein the step of placing a skin patch on the expanded foam comprises laying up a plurality of uncured composite layers on the expanded foam.
 8. The method according to claim 1, wherein the skin patch is of a composite material.
 9. The method according to claim 1, wherein the step of placing a skin patch on the expanded foam comprises laying up a composite layers across the expanded foam and an adjacent skin surface.
 10. The method according to claim 1, further comprising: curing the skin patch.
 11. The method according to claim 10, wherein the step curing the skin patch includes soaking the skin patch at an ambient temperature.
 12. The method according to claim 10, wherein the step curing the skin patch creates a bond between the expanded foam and an inner surface of the skin patch.
 13. The method according to claim 10, wherein the step curing the skin patch includes allowing a resin in the skin patch forming a structural bond between the expanded foam and the skin patch.
 14. The method according to claim 10, wherein the step curing the skin patch does not require an artificial heat source.
 15. A method of repairing a core stiffened structure having a first skin, a core member, and a second skin, the method comprising: removing a damaged portion of the first skin and the core member; preparing a foam mixture; pouring the foam mixture into a cavity, the cavity being created by the step of removing the damaged portion of the first skin and the core member; allowing the foam mixture to expand and fill the cavity so as to form an expanded foam; placing a skin patch on the expanded foam.
 16. The method according to claim 15, wherein the step of preparing the foam mixture includes mixing a first resin component with a second resin component so as to form a homogenious mixture.
 17. The method according to claim 15, further comprising: restricting the expansion of the foam mixture with a mold tool, the mold tool having an interior surface that resembles the desired outer contour of the expanded foam.
 18. The method according to claim 15, further comprising: cleaning the cavity prior to pouring the foam mixture into the cavity.
 19. The method according to claim 15, wherein the step of placing the skin patch on the expanded foam includes laying up a plurality of composite layers over the skin patch.
 20. The method according to claim 19, wherein the composite layers are uncured pre-preg composite material. 